Small gas turbine

ABSTRACT

A small gas turbine, in particular for powering airplane models, provided with a casing (3), a compressor rotor (1, 2), a combustion chamber (15-17) and a turbine rotor (5), the turbine rotor (5) driving the compressor rotor(1, 2) by a shaft positioned in a shaft tunnel extending through the combustion chamber (15-17) and surrounded by a helical fuel feed and vaporizing tube (22, 23), the combustion chamber (15-17) being defined by a cylindrical outer wall (15) and a frusto-conical inner wall (16) provided with combustion air intake orifices (19-21)the outlet side of the inner wall (16) of the combustion chamber extending into a cylindrical flow element (6, 7) forming an annular gap therewith, which in turn extends into the center opening of the turbine side guide vane ring (4) and is supported by the end (9) of the shaft tunnel (8), and the other end of the shaft tunnel (8) being secured to the casing (3) between the compressor rotor (1, 2) and the front wall (17) of the combustion chamber.

The invention relates to a small gas turbine, in particular forpropelling model aircraft, provided with a rotationally symmetricalcasing, a compressor rotor at the fluid intake of the casing, acombustion chamber arranged within the casing and a turbine rotor with astationary guide vane ring arranged forwardly thereof at the fluidoutlet of the casing, the turbine rotor driving the compressor rotor byway of a shaft provided in a shaft tunnel extending through thecombustion chamber, the tunnel being surrounded by a helical fuel feedand vaporizing tube arranged within the combustion chamber and feedinginto it by means of nozzles.

A small gas turbine of this kind has been repeatedly disclosed by thepresent inventor: FMT Flug- u. Modelltechnik 405-10/89 pages 20/21; FMTFlug- und Modelltechnik 408-1/90 pages 52/53; FMT Flug- u. ModelltechnikSpezial Scale 89/90 pages 69/71; FMT Flug- u. Modelltechnik SpezialScale Nr.1 90 pages 34/37 (all journals are published by Verlag fuerTechnik und Handwerk, Baden-Baden, Germany). In accordance with thelatest publications, the gas turbine designated "Strahlturbine FD3"(turbojet FD3) possesses the following technical specifications: Thefuel used in flight operations is diesel fuel or a mixture of dieselfuel and gasoline. In a stationary operation, propane or butane, ingaseous or liquid state, may be used appropriately metered. The radial(centrifugal) compressor stage is provided with a rotor with backwardlycurved blades and a cover plate. The turbine rotor is acting axially.Fuel is vaporized in a helical tube which also acts as a cooling coilfor the shaft tunnel.

However, this state of the art still suffers from the following defects.In small gas turbines, the combustion of liquid fuel in a relativelysmall combustion chamber is especially problematical. Comparably sizedsmall gas turbines, for instance for starting large turbo jet engines,preferably use reverse flow annular combustors or disc-shapedcombustors. Either structure would, however, detrimentally increase massand volume of a small gas turbine, particularly of the kind used forpropelling model aircraft. An optimum combustor structure yieldinguniform combustion with a stabilized flame front even under rapidlychanging loads, has not yet been devised.

At the same time, there is the problem that the shaft tunnel in one wayor another must extend through the combustion chamber, in order to driveof the antecedent compressor stage. This leads to thermal problems: Incase of insufficient heat dissipation, the shaft tunnel supported by theguide vane ring at the turbine side, because of unavoidable temperaturedifferences, causes deformation of the guide vane ring and, hence, anoff-center displacement of the turbine rotor. To avoid any grindingcontact, a relatively large free play of the turbine must be providedwithin the casing outlet, which results in a reduced internal efficiencyof the turbine as well as the thrust/weight ratio. Furthermore,over-heating of the shaft tunnel often leads to a malfunction of thebearing of the turbine rotor supported in the outlet end of the wavetunnel and thus to a significantly reduced uniformity in the rotationsof the turbine rotor as well.

Therefore, an optimum structure of the combustion chamber and itssupport must also take these thermal problems into consideration.

It is an object of the invention to provide a small gas turbine of thekind referred to at the outset, in which uniform stabilized combustionis attained even under rapid load changes and in which combustion heatis dissipated in a controlled manner.

Further objects of the invention reside in providing a gas turbine whichfor propelling model aircraft may preferably operated with diesel fuel;which in a stationary operation may optionally also be driven withgaseous fuels, preferably propane and/or butane gas, and which requiresno adaptations or changes in the gas turbine when the kind of the fuelis changed; which attains a maximum thrust of about 30N at athrust/weight ratio of at least 3; and which may be commerciallymanufactured.

In accordance with the invention these objects are accomplished in a gasturbine of the kind referred to in the introduction, by the combustionchamber being defined by a cylindrical outer wall and a frusto-conicalinner wall surrounding the shaft tunnel and provided with combustion airintake orifices, the conical base of which faces the compressor rotorand is connected to the outer wall of the combustion chamber by way ofan annular front wall, with the inner wall at the outlet side end of thecombustion chamber extending, forming an annular gap, into a cylindricalflow element which in turn is positioned, forming an annular gap, in thecenter opening of the guide vane ring at the outlet and is supported bythe end of the shaft tunnel, the other end of the shaft tunnel beingsecured to the casing between the compressor rotor and the front wall ofthe combustion chamber.

This construction and mounting of the combustion chamber ensures auniform combustion with a stabilized flame front, even at rapid loadchanges when changing the fuel supply, as well as a thermal disjunctionbetween the guide vane ring at the outlet side and the turbine rotorbearing.

It has been found that the-preferred diameter of the inner wall of thecombustion chamber at the base of the cone is 0.7 times the diameter ofthe outer wall of the combustion chamber, and that the width of theannular gap preferably is about 0.05 to 0.1 mm.

In order to mount the shaft tunnel in a vibration-free manner in thistype of combustion chamber support in a vibration-safe way it isadvantageous to secure the intake side end of the shaft tunnel to thecasing by at least three braces radially extending between thecompressor rotor and the front wall of the combustion chamber. In thatcase the braces may be affixed to the casing by radial bolts which alsosecure a cover radially extending over the compressor rotor but leavingthe center intake section free, to provide for a gas turbine of simpleconstruction.

The assembly of the small gas turbine is further simplified by thebraces supporting a guide vane ring positioned at the intake between thecompressor rotor and the front wall of the combustion chamber, such aring being commonly provided in such turbines. The guide vane ring mayat the same time be used to define the position of the shaft tunnel byan inner detent in the center opening of the guide vane ring at itsintake side to provide a shoulder for securing the shaft tunnel.

Commercial manufacture is further facilitated by pressing the combustionchamber at the outlet side of its outer wall against the inner wall ofthe casing tapering in the fluid outlet direction, with compressionsprings acting between the intake side guide vane ring and the frontwall of the combustion chamber.

In accordance with a preferred embodiment of the invention the mountingof the flow element at the end of the shaft tunnel provides for asubstantially annular air flow gap aligned with an annular gap betweenthe shaft tunnel and the outlet end of the inner wall of the combustionchamber. In this manner the inside of the inner wall of the combustionchamber serves at the same time to conduct cooling air to the outletside section of the shaft tunnel, to the turbine rotor bearing, and tothe center portion of the turbine rotor. Cooling of these elements isthus significantly amplified.

It has been found that the width of this air flow gap preferably shouldbe about 0.25-0.5 mm.

In accordance with a further embodiment of the invention the hot end ofthe fuel feed and vaporizing tube is bent back toward the front wall ofthe combustion chamber and extends along the marginal section betweenthe front wall and the inner wall of the combustion chamber as anannular manifold provided with nozzles. This provides for a uniformradially symmetric temperature distribution within the combustionchamber and, hence, for a uniform impact on the turbine stage, which isindispensable for a reliable operation. Preferably, this marginalsection is, in cross-section, semi-circularly rounded off. Placing themanifold into a different position within the combustion chamber wouldresult in an asymmetric temperature distribution at the annular manifoldbecause of the flow of hot gas and the simultaneous emission ofvaporized fuel from the nozzles, and thus in an asymmetric emission ofthe vaporized fuel which in turn leads to a non-uniform temperaturedistribution within the turbine stage.

The helical section of the fuel feed and vaporizing tube may bemanufactured in a cost-efficient manner from one piece of tubing of heatresistant material. Preferably, the cross-sectional dimension of thefuel feed and vaporizing tube is periodically constricted at intervalsof 1-2 cm, or a metal ball chain may be inserted into the fuel feed andvaporizing tube, with the diameter of the balls of the chain beingsomewhat less than the inner diameter of the fuel feed and vaporizingtube. Either measure improves heat transfer from the fuel feed andvaporizing tube to the fuel flowing therein, because the laminar flowwithin the tube is forced into turbulence.

A further advantageous embodiment of the invention is characterized bythe front wall of the combustion chamber being provided with radial airintake slots disposed angularly relative to the axis of the casing. Thisresults in flow turbulence within the combustion chamber, with arotational axis the same as the axis of the casing, thus improving theuniform distribution and combustion of the vaporized fuel within thecombustion chamber. For the same reason, it is advantageous to bend themargins of those combustion air intake orifices of the inner wall of thecombustion chamber which are positioned close to the base of the coneaway from the surface of the cone, so that the direction of these intakeorifices extends at an angle with respect to the radial direction.

Preferably, the outer wall of the combustion chamber is provided withadjustable air flaps to allow fine tuning of the temperature gradient ina radial direction at the turbine stage.

Advantageously, the edge of the outer wall of the combustion chamber atthe outlet side supported by the inner wall of the casing is alsoprovided with air slots to provide for air flow to the rear section ofthe combustion chamber.

In the small gas turbine in accordance with the invention it isparticularly advantageous to utilize as the compressor rotor a radially(centrifugally) compressing rotor provided with a cover ring and rotorblades deflected in the outlet direction. This results in a compressorstage of very high internal efficiency at a minimum effort as regardsthe guide system. The axial play of a covered compressor rotor isrelatively uncritical. The backward deflection of the blades results ina compressor rotor which during start-up acts as a turbine, so that avery small blower with less than 20 W output power is sufficient tostart the small gas turbine.

Preferably, the blade height at the intake of the compressor rotor is atleast 1.5 times the vane height at the outlet. The invention willhereinafter be described in detail with reference to an embodimentdepicted in the accompanying drawings. In the drawings:

FIG. 1 depicts an axial section of a small gas turbine in accordancewith the invention;

FIGS. 2 and 3 schematically depict the structure of the jets in theannular manifold;

FIGS. 4 and 5 show the structure of the air flaps in the outer wall ofthe combustion chamber;

FIGS. 6 and 7 show the structure of the combustion air intake orificesin the inner wall of the combustion chamber at the intake side;

FIGS. 8 and 9 show the structure of the air slots in the front wall ofthe combustion chamber; and

FIG. 10 depicts the constrictions in the fuel feed and vaporizing tube.

The small gas turbine depicted in the drawings is particularly suitedfor powering model aircraft and comprises a rotationally symmetriccasing 3, a compressor rotor 1, 2 in the fluid intake of the casing acombustion chamber 15-17 arranged within the casing, and a turbine rotor5 in the fluid outlet of the casing. A stationary guide vanes ring 4 ispositioned in front of the turbine rotor 5.

The turbine rotor 5 drives the compressor rotor 1, 2 by means of a shaftwhich is positioned in a shaft tunnel 8 extending through the combustionchamber 15-17. The shaft tunnel 8 is surrounded by a helical fuel feedand vaporizing tube 22, 23 arranged within the combustion chamber andopening into it by way of nozzles 24.

The combustion chamber 15-17 is defined by a cylindrical outer wall 15,a frusto-conical inner wall 16, and a circular front wall 17.

The inner wall 16 of the combustion chamber surrounds the shaft tunnel 8and is provided with combustion air intake orifices 19-21. The base ofthe cone of the inner wall 16 of the combustion chamber faces thecompressor rotor 1, 2 and is connected with the outer wall 15 of thecombustion chamber by means of the front wall 17 of the combustionchamber.

The outlet side end of the inner wall 16 of the combustion chamberextends into a cylindrical flow element 6, 7 forming an annular gapbetween them. The cylindrical flow element 6, 7 is in turn positioned inthe center opening of the circular vane ring 4, forming an annular gap,and is supported by the shaft tunnel 8.

The other end of the shaft tunnel 8 secured to the casing 3 between thecompressor rotor 1, 2 and the front wall 17 of the combustion chamber.

The diameter of the inner wall 16 of the combustion chamber at the baseof the cone is 0.7 times the diameter of the outer wall 15 of thecombustion chamber. The width of the annular gap between the flowelement 6, 7 and the guide van ring 4 at the turbine side is about0.05-0.1 mm.

The end of the shaft tunnel 8 at the intake side is secured to thecasing 3 by at least three braces 10 which radially extend between thecompressor rotor 1, 2 and the front wall 17 of the combustion chamber.The braces are fastened to the casing by radial bolts 11 which alsosecure a cover 12 extending over the compressor rotor 1, 2 but leavingits fluid intake center open. The braces 10 additionally support a guidevane ring 13 arranged at the intake side between the compressor rotor 1,2 and the front wall 17 of the combustion chamber.

The center opening of the guide vane ring 13 at the intake side isprovided with an inner detent which provides a shoulder for axiallysecuring the shaft tunnel 8.

By way of its outer wall 15 at the outlet side the combustion chamber15-17 is pressed against the inner wall of the casing 3 which narrows inthe direction of fluid outlet, by at least two compression springs 18which act between the guide vane ring 13 at the intake and the frontwall 17 of the combustion chamber.

The mounting of the flow element 6, 7 at the end of the shaft tunnel 8is such that a substantially annular air flow gap results which isaligned with an annular gap between the shaft tunnel 8 and the end ofthe inner wall 16 of the combustion chamber at the outlet side. Thewidth of the air flow gap between the flow element 6, 7 and the shafttunnel 8 is about 0.25-0.5 mm.

The hot end of the fuel feed and vaporizing tube 22, 23 is bent back tothe front wall 17 of the combustion chamber and extends along themarginal section between the front wall 17 and the inner wall 16 of thecombustion chamber as an annular manifold 23 provided with nozzles 24.In cross section, the marginal section is of semicircular configuration.

The cross-section of the fuel feed and vaporizing tube 23, 24 isnarrowed by constrictions 26 at intervals of 1-2 cm. Alternatively, ametal ball chain may be inserted into the fuel feed and vaporizing tube22, 23, the diameter of the balls of which is somewhat smaller than theinternal diameter of the fuel feed and vaporizing tube.

The front wall 17 of the combustion chamber is provided with radial airintake slots 25 the longitudinal direction of which extends angularlyrelative to the axis of the casing.

The margins of those combustion air intake orifices 19 of the inner wall16 of combustion chamber which are located near the base of the cone,have been pried out of the surface of the cone so that these intakeorifices are directed at an angle relative to the radial direction.

The outer wall 15 of the combustion chamber is further provided withadjustable air flaps. The flaps are adjusted by bending more or lesspreferably in the direction of the combustion chamber.

The front edge of the outer wall 15 of the combustion chamber supportedby the inner wall of the casing is provided with air slots 28.

The compressor rotor 1, 2 compresses in a radial direction, is providedwith a cover ring 1 and its blades 2 are deflected in the outletdirection. The height of the blades at the intake of the compressorrotor 1, 2 is at least 1.5 times larger than the height of the blades atthe outlet.

What is claimed is:
 1. Small gas turbine, in particular for poweringmodel aircraft, comprising a rotationally symmetric casing (3), acompressor rotor (1, 2) at the fluid intake of the casing, a combustionchamber (15-17) arranged within the casing and a turbine rotor (5) inthe fluid outlet of the casing provided with a stationary guide vanering (4) positioned in front thereof, the turbine rotor (5) driving thecompressor rotor (1, 2) by a shaft provided in a shaft tunnel (8)extending through the combustion chamber (15-17), the shaft tunnel beingsurrounded by a helical fuel feed and vaporizing tube (22, 23) arrangedwithin the combustion chamber and communicating with it by nozzles (24),characterized bythe combustion chamber (15-17) being defined by acylindrical outer wall (15) and a frusto-conical inner wall (16)surrounding the shaft tunnel (8) and provided with combustion air intakeorifices (19-21), the base of the cone of which faces the compressorrotor (1, 2) and is connected with the outer wall (15) of the combustionchamber by a circular front wall (17), the outlet side end of the innerwall (16) of the combustion chamber extending into a cylindrical flowelement (6, 7) forming an annular gap therewith, which element is inturn positioned within the center opening of the outlet side guide vanering (4) forming an annular gap therewith and is supported by the end ofthe shaft tunnel (8), and the other end of the shaft tunnel (8) beingsecured to the casing between the compressor rotor (1, 2) and the frontwall (17) of the combustion chamber.
 2. Small gas turbine in accordancewith claim 1, characterized by the diameter of the inner wall (16) ofthe combustion chamber at the base of the cone being 0.7 times thediameter of the outer wall (15) of the combustion chamber.
 3. Small gasturbine in accordance with claim 1, characterized by the width of theannular gap between the flow element (6, 7) and the guide vane ring (4)at the turbine side is about 0.05-0.1 mm.
 4. Small gas turbine inaccordance with claim 1, characterized by the intake side end of theshaft tunnel (8) being secured to the casing (3) by at least threebraces (10) radially extending between the compressor rotor (1, 2) andthe front wall (17) of the combustion chamber.
 5. Small gas turbine inaccordance with claim 4, characterized by the braces (10) being fastenedto the casing (3) by radial bolts (11) which also secure a cover (12)extending over the compressor rotor (1, 2) and leaving its fluid intakecenter open.
 6. Small gas turbine in accordance with claim 4,characterized by the braces (10) supporting a guide vane ring (13) atthe intake side between the compressor rotor (1, 2) and the front wall(17) o the combustion chamber.
 7. Small gas turbine in accordance withclaim 6, characterized by the center opening of the intake side guidevane ring (13) being provided with an inner detent providing a shoulderfor axially securing the shaft tunnel (8).
 8. Small gas turbine inaccordance with claim 6, characterized by the combustion chamber (15-17)being pressed by the end of its outer wall (15) at the outlet sideagainst the inner wall of the casing (3) narrowing in the direction offluid outlet by compression springs (18) acting between the intake sideguide vane ring (13) and the front wall (17) of the combustion chamber.9. Small gas turbine in accordance with claim 1, characterized by themounting of the flow element (6, 7) at the end of the shaft tunnel (8)providing a substantially annular air flow gap aligned with an annulargap between the shaft tunnel (8) and the outlet side end of the innerwall (16) of the combustion chamber.
 10. Small gas turbine in accordancewith claim 9, characterized by the width of the air flow gap between theflow element (6, 7) and the shaft tunnel (8) being about 0.25-0.5 mm.11. Small gas turbine in accordance with claim 1, characterized by thehot end of the fuel feed and vaporizing tube (22, 23) being bent backtoward the front wall (17) of the combustion chamber and extends alongthe marginal section between the front wall (17) and the inner wall (16)of the combustion chamber as an annular manifold 23 provided withnozzles (24).
 12. Small gas turbine in accordance with claim 11,characterized by the marginal section being in cross-sectionsemi-circularly rounded off.
 13. Small gas turbine in accordance withclaim 1, characterized by the cross-section of the fuel feed andvaporizing tube (23, 24) being narrowed by constrictions (26) atintervals of 1-2 cm.
 14. Small gas turbine in accordance with claim 1,characterized by a metal ball chain being inserted into the fuel feedand vaporizing tube (22, 23), the diameter of the balls of which isslightly less than the internal diameter of the fuel feed and vaporizingtube.
 15. Small gas turbine in accordance with claim 1, characterized bythe front wall (17) of the combustion chamber being provided with radialair intake slots (25) the longitudinal direction of which extendsangularly relative to the axis of the casing.
 16. Small gas turbine inaccordance with claim 1, characterized by the margins of the combustionair intake orifices (19) of the inner wall (16) of the combustionchamber are located near the base of the cone, are pried out of thesurface of the cone so that the direction of the intake orifices (19)extends angularly relative to the radial direction.
 17. Small gasturbine in accordance with claim 1, characterized by the outer wall (15)of the combustion chamber being provided with adjustable air flaps (27).18. Small gas turbine in accordance with claim 1, characterized by theoutlet side front edge of the outer wall (15) of the combustion chambersupported at the inner wall of the casing (3) being provided with airslots (28).
 19. Small gas turbine in accordance with claim 1,characterized by the compressor rotor (1, 2) compressing in a radialdirection, being provided with an annular cover (1) and that its blades(2) are deflected in the outlet direction.
 20. Small gas turbine inaccordance with claim 19, characterized by the height of the blades atthe intake of the compressor rotor (1, 2) being at least 1.5 times theheight of the blades at the outlet.